CN109583041A - A kind of Craft Orbit design method - Google Patents

A kind of Craft Orbit design method Download PDF

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CN109583041A
CN109583041A CN201811321325.4A CN201811321325A CN109583041A CN 109583041 A CN109583041 A CN 109583041A CN 201811321325 A CN201811321325 A CN 201811321325A CN 109583041 A CN109583041 A CN 109583041A
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boosting
last
burnout
orbit
thresholding
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CN109583041B (en
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梁纪秋
刘萧磊
张天翼
陈腾芳
胡长伟
黄雷
彭威
韩通
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General Designing Institute of Hubei Space Technology Academy
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    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
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    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
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Abstract

The invention discloses a kind of Craft Orbit design methods, are related to Craft Orbit design field, including initial trajectory is designed and generated according to launch mission, obtain conditional parameter of entering the orbit；Flight status parameter is calculated in the speed of launch inertial coordinate system and position according to carrier rocket；Initial calculation condition is set, using last boosting first time Burnout altitude of the apogee, last second of shutdown point speed of boosting and local trajectory tilt angle and target track height as constraint condition, according to initial calculation condition, enter the orbit conditional parameter and flight status parameter is calculated, obtain the deviation of constraint condition, initial calculation condition is optimized, Optimal calculation condition is obtained, calculates simultaneously outputting standard ballistic data accordingly.Craft Orbit design method of the invention is reached the design object of optimization inter-orbital transfer time, is provided Optimal calculation condition for ballistic computation based on energy management, effectively reduces the Orbit Transformation working time, and realization is quickly entered the orbit.

Description

A kind of Craft Orbit design method
Technical field
The present invention relates to Craft Orbit design fields, and in particular to a kind of Craft Orbit design method.
Background technique
The more star transmittings of carrier rocket, also known as several satellite in a rocket, refer to that multi-satellite is emitted to pre- orbit determination by one piece of carrier rocket Road.Several satellite in a rocket radiation pattern can fully utilize the carrying capacity of carrier rocket, reduce the launch cost of carrier rocket, more stars Transmitting is exactly to be provided with multi-satellite in radome fairing, and the satellite in radome fairing carries out reasonable arrangement according to how many and size, according to Launch mission actual demand, the height certain to space are launched together or separately.
In launch mission, if needing to deliver when discharging different types of satellite load in different orbit altitudes Rocket carries out Orbit Transformation, based on traditional Orbital Transfer, is generally completed using classical Huo Man transfer, Huo Man transfer Though in practical implementation, since the time of Huo Man transfer is longer, giving each point of carrier rocket using minimal consumption fuel as criterion System work proposes higher reliability and stability requirement, during completing launch mission, increases to launch mission Risk.
Summary of the invention
In view of the deficiencies in the prior art, the purpose of the present invention is to provide a kind of Craft Orbit design sides Method realizes that carrier rocket is quickly entered the orbit.
To achieve the above objectives, the technical solution adopted by the present invention is that: a kind of Craft Orbit design method, packet It includes:
Initial trajectory is designed and generated according to launch mission, obtains conditional parameter of entering the orbit；
Flight status parameter is calculated in the speed of launch inertial coordinate system and position according to carrier rocket；
Initial calculation condition is set, is shut down for the second time spot speed with last boosting first time Burnout altitude of the apogee, last boosting Degree and local trajectory tilt angle and target track height are constraint condition, and according to initial calculation condition, entering the orbit conditional parameter and flies Row state parameter is calculated, and the deviation of constraint condition is obtained, and is optimized to initial calculation condition, and Optimal calculation item is obtained Part calculates simultaneously outputting standard ballistic data accordingly.
Based on the above technical solution, initial calculation condition includes last boosting first time Burnout altitude of the apogee Second RE1, last boosting first time shutdown residual propellant quality DSMY41, last boosting of residual propellant quality of shutting down The attitude angle PHI4 of DSMY42, end second of starting point locality trajectory tilt angle FIRE1 of boosting and the first task of last boosting.
Based on the above technical solution, initial calculation condition is optimized, obtains Optimal calculation condition, it is specific to wrap It includes:
Four adjustment thresholdings are set；
Last boosting shutdown residual propellant quality DSMY41 for the first time is adjusted, keeps last boosting first time Burnout apogee high Inclined absolute value of the difference is spent in the first thresholding；
Second of residual propellant quality DSMY42 that shuts down of last boosting is adjusted, second of shutdown point speed deviation of last boosting is made Absolute value in the second thresholding；Also, last second of starting point locality trajectory tilt angle FIRE1 of boosting is adjusted, last boosting the is made The absolute value of secondary Burnout locality ballistic inclination deviation is in third thresholding；
Last boosting first time Burnout altitude of the apogee RE1 is adjusted, keeps second of Burnout track major semiaxis of last boosting inclined Absolute value of the difference is in the 4th thresholding.
Based on the above technical solution, flight status parameter includes injection point carrier rocket to the earth's core actual range The practical absolute velocity V of RE, the injection point carrier rocket and practical local trajectory tilt angle η of injection point carrier rocket, and
Wherein, speed parameter of the carrier rocket in launch inertial coordinate system is VIxg, VIyg, VIzg, and location parameter is RExg,REyg,REzg。
Based on the above technical solution, conditional parameter of entering the orbit includes injection point carrier rocket to the earth's core theoretical distance BzRE, injection point carrier rocket theory absolute velocity VbzAnd the theoretical local trajectory tilt angle η of injection point carrier rocketbz。
Based on the above technical solution, the calculation formula of last boosting first time Burnout altitude of the apogee RE1 is as follows:
RE1=bzRE-R0
Wherein, R0For earth radius.
Based on the above technical solution, last boosting first time Burnout altitude of the apogee deviation delta h1 are as follows:
Δ h1=h_apo-RE1
Wherein, h_apo is last boosting first time Burnout apogee computed altitude；
If the last inclined absolute value of the difference of boosting first time Burnout altitude of the apogee exceeds the first thresholding, last boosting is for the first time The adjustment amount of shutdown residual propellant quality DSMY41 is Δ h/k1, k1For the first iteration coefficient.
Based on the above technical solution, second of shutdown point speed deviation delta V of last boosting are as follows:
Δ V=V-Vbz
Above-mentioned end second of Burnout locality ballistic inclination deviation Δ η of boosting are as follows:
Δ η=η-ηbz
If the last inclined absolute value of the difference of second of shutdown point speed of boosting exceeds the second thresholding of thresholding, second of residue of shutting down The adjustment amount of propellant mass DSMY42 is Δ V/k2, k2For secondary iteration coefficient；If last second of Burnout of boosting trajectory incline The absolute value of angular displacement exceeds third thresholding, then the adjustment amount of second of starting point locality trajectory tilt angle FIRE1 of last boosting is Δ η。
Based on the above technical solution, second of Burnout track major semiaxis deviation delta h2 of last boosting are as follows:
Δ h2=a-bzRE
Wherein, a is second of Burnout track major semiaxis of last boosting；
If last second of inclined absolute value of the difference of Burnout track major semiaxis of boosting exceeds the 4th thresholding, last boosting is for the first time The adjustment amount of Burnout altitude of the apogee RE1 is Δ h2.
Based on the above technical solution, the first thresholding e1For 50m, above-mentioned second thresholding e2For 0.01m/s, above-mentioned Three thresholding e3It is 0.001 °, above-mentioned 4th thresholding e4For 100m.
Compared with the prior art, the advantages of the present invention are as follows:
Craft Orbit design method of the invention, comprehensively considers transfer time and energy consumption, from different tracks When being shifted in height, by the energy consumption of flexible allocation carrier rocket final stage, based on energy management, reach optimization The design object of inter-orbital transfer time provides Optimal calculation condition for ballistic computation, effectively reduces the Orbit Transformation working time, real Now quickly enter the orbit.
Detailed description of the invention
Fig. 1 is the flow chart of Craft Orbit design method in the embodiment of the present invention.
Specific embodiment
Invention is further described in detail with reference to the accompanying drawings and embodiments.
Shown in Figure 1, the embodiment of the present invention provides a kind of Craft Orbit design method comprising:
Initial trajectory is designed and generated according to launch mission, obtains conditional parameter of entering the orbit；
Flight status parameter is calculated in the speed of launch inertial coordinate system and position according to carrier rocket；
Initial calculation condition is set, is shut down for the second time spot speed with last boosting first time Burnout altitude of the apogee, last boosting Degree and local trajectory tilt angle and target track height are constraint condition, and according to initial calculation condition, entering the orbit conditional parameter and flies Row state parameter is calculated, and the deviation of constraint condition is obtained, and is optimized to initial calculation condition, and Optimal calculation item is obtained Part calculates simultaneously outputting standard ballistic data accordingly.
The embodiment of the present invention comprehensively considers transfer time and energy consumption, when being shifted from different orbit altitudes, By the energy consumption of flexible allocation carrier rocket final stage, based on energy management, reach setting for optimization inter-orbital transfer time Target is counted, Optimal calculation condition is provided for ballistic computation, effectively reduces the Orbit Transformation working time, reach the mesh quickly entered the orbit 's.
The embodiment of the present invention uses Three Degree Of Freedom model trajectory to calculate trajectory at the standard conditions, and standard conditions include: ground Spherical model is IAG-75 ellipsoidal model；The acceleration of gravity of earth surface is 9.80665m/s2；Atmospheric conditions are that national standard is big Gas；Also, carrier rocket final stage works twice, and load is carried out orbit maneuver from the preliminary orbit height of launch mission and is sent into mesh Mark orbit altitude, wherein preliminary orbit is 500km circular orbit, and target track height is 700km circular orbit.
Preferably, initial calculation condition includes last boosting first time Burnout altitude of the apogee RE1, last boosting pass for the first time Machine residual propellant quality DSMY41, it second of residual propellant quality DSMY42 that shuts down of last boosting, is switched on for second of last boosting The attitude angle PHI4 of the first task of point locality trajectory tilt angle FIRE1 and last boosting.
Preferably, above-mentioned that initial calculation condition is optimized, Optimal calculation condition is obtained, is specifically included:
Four adjustment thresholdings, including the first thresholding, the second thresholding, third thresholding and the 4th thresholding are set, wherein first Limit e1For last boosting first time Burnout altitude of the apogee deviation iteration threshold, the second thresholding e2For second of Burnout of last boosting Velocity deviation iteration threshold, third thresholding e3For second of Burnout locality ballistic inclination deviation iteration threshold of last boosting, the 4th Thresholding e4For second of Burnout track major semiaxis deviation iteration threshold of last boosting；
The first task of carrier rocket final stage terminates, and adjusts last boosting shutdown residual propellant quality DSMY41 for the first time, Make the inclined absolute value of the difference of last boosting first time Burnout altitude of the apogee in the first thresholding；
After carrier rocket final stage coasting-flight phase, the second task of carrier rocket final stage terminates, and adjusts second of last boosting Shut down residual propellant quality DSMY42, makes the inclined absolute value of the difference of second of shutdown point speed of last boosting in the second thresholding；And And second of starting point locality trajectory tilt angle FIRE1 of last boosting is adjusted, keep second of Burnout locality trajectory tilt angle of last boosting inclined Absolute value of the difference is in third thresholding；
Last boosting first time Burnout altitude of the apogee RE1 is adjusted, keeps second of Burnout track major semiaxis of last boosting inclined Absolute value of the difference is in the 4th thresholding.
In the embodiment of the present invention, carrier rocket launch inertial coordinate system o is definedAXyz, coordinate origin and carrier rocket emit Point is connected, oAX-axis is directed toward transmitting sighted direction, o in launch point horizontal planeAY-axis is directed toward perpendicular to launch point horizontal plane face Side, oAZ-axis is perpendicular to xoAThe face y simultaneously constitutes right-handed coordinate system.After rocket takes off, point and each axis direction of coordinate system are protected in inertial space It holds constant.Certain moment, velocity parameter values of the carrier rocket in launch inertial coordinate system are VIxg, VIyg, VIzg, location parameter Value is RExg, REyg, REzg.
Wherein, flight status parameter includes that injection point carrier rocket is real to the earth's core actual range RE, injection point carrier rocket The border absolute velocity V and practical local trajectory tilt angle η of injection point carrier rocket.According to carrier rocket in launch inertial coordinate system Speed and location parameter are calculated:
In the embodiment of the present invention, conditional parameter of entering the orbit include injection point carrier rocket to the earth's core theoretical distance bzRE, enter the orbit Point carrier rocket theory absolute velocity VbzAnd the theoretical local trajectory tilt angle η of injection point carrier rocketbz。
Wherein, injection point carrier rocket is designed to the earth's core theoretical distance bzRE according to launch mission demand, injection point Carrier rocket absolute velocity VbzIt is calculated according to injection point carrier rocket to the earth's core theoretical distance bzRE, injection point carrier rocket Theoretical locality trajectory tilt angle ηbzGenerally take 90 °.
In the embodiment of the present invention, ballistic computation needs to be arranged initial calculation condition, and initial calculation condition need to design rationally, The case where avoiding track that from can not iterating to calculate.
Preferably, the calculation formula of last boosting first time Burnout altitude of the apogee RE1 is as follows:
RE1=bzRE-R0
Wherein, R0For earth radius, R0=6378140m.
Second of shutdown residual propellant quality of last boosting first time shutdown residual propellant quality DSMY41 and last boosting DSMY42 is configured according to different carrier rocket model population parameters, and DSMY41 > DSMY42.Last the first task of boosting The range that sets of attitude angle PHI4 size as 0 ° < PHI4 < 180 °.
In the present embodiment, each iteration coefficient need to meet a certain range requirement, if iteration coefficient is excessive, iteration speed compared with Slowly, if coefficient is too small, given target track requirement is unable to satisfy after possible iteration.Specifically, last boosting is closed for the first time Machine point altitude of the apogee deviation delta h1 are as follows:
Δ h1=h_apo-RE1
Wherein, h_apo is last boosting first time Burnout altitude of the apogee.
It is used to define celestial sphere reference frame in the celestial equator and equinox at J2000 moment, the reference frame is also writeable Make J2000 coordinate or is simply denoted as J2000.
Specifically, the velocity parameter values according to carrier rocket in launch inertial coordinate system are VIxg, VIyg, VIzg, position Setting parameter value is RExg, REyg, REzg, and the speed V for obtaining J2000 system is converted by coordinateJ2000,Vj2000,VJ2000With position RJ2000,Rj200,0RJ2000Size, by the velocity magnitude V of J2000 systemJ2000,Vj2000,VJ2000It is defined asBy the position of J2000 system Size RJ2000,Rj2000,RJ2000It is defined as r, steps are as follows for the calculating of h_apo:
(1) moment of momentum h and moment of momentum modulus h size of carrier rocket are calculated:
(2) basisThe size of r and h calculates e and e:
(3) basis | h |, | e | calculate h_apo size:
H_apo=a (1+ | e |)
Wherein, μ is Gravitational coefficient of the Earth, μ=3.986005 × 1014；A is that last second of Burnout track of boosting is long by half Axis.
If the last inclined absolute value of the difference of boosting first time Burnout altitude of the apogee exceeds the first thresholding e1, i.e., | Δ h | > e1, Then the adjustment amount of last boosting first time shutdown residual propellant quality DSMY41 is Δ h/k1, k1For the first iteration coefficient, and 1e3 ≤k1≤1e4；If | Δ h |≤e1, then iteration terminates, that is, meets last boosting first time Burnout altitude of the apogee constraint.
Specifically, second of shutdown point speed deviation delta V of last boosting are as follows:
Δ V=V-Vbz
Second of Burnout locality ballistic inclination deviation Δ η of last boosting are as follows:
Δ η=η-ηbz
If the last inclined absolute value of the difference of second of shutdown point speed of boosting exceeds the second thresholding of thresholding e2, i.e., | Δ V | > e2, then The adjustment amount of second of shutdown residual propellant quality DSMY42 is Δ V/k2, k2For secondary iteration coefficient, and 1≤k2≤20；If Second of Burnout of last boosting ballistic inclination deviation absolute value exceed third thresholding e3, i.e., | Δ η | > e3, then last boosting the The adjustment amount of secondary starting point locality trajectory tilt angle FIRE1 is Δ η；If | Δ V |≤e2And | Δ η |≤e3, then iteration terminates, i.e., Meet last second of shutdown point speed of boosting and local trajectory tilt angle constraint.
Specifically, second of Burnout track major semiaxis deviation delta h2 of last boosting are as follows:
Δ h2=a-bzRE
Wherein, a is second of Burnout track major semiaxis of last boosting；A is according to carrier rocket end boosting as described above Coordinated Universal Time(UTC), that is, UTC time of secondary shutdown point moment, the speed under J2000 coordinate system, position size are calculated.
If last second of inclined absolute value of the difference of Burnout track major semiaxis of boosting exceeds the 4th thresholding e4, i.e., | Δ h | > e4, Then the adjustment amount of last boosting first time Burnout altitude of the apogee RE1 is Δ h2；If | Δ h |≤e4, then iteration terminates, that is, meets Target track is highly constrained.
Preferably, the first thresholding e1For 50m, the second thresholding e2For 0.01m/s, third thresholding e3It is 0.001 °, the 4th thresholding e4For 100m.
The embodiment of the present invention is suitable for that load is sent into target track height from preliminary orbit height in carrier rocket final stage Change rail situation, guarantee carrier rocket accurately enter the orbit.
The present invention is not limited to the above-described embodiments, for those skilled in the art, is not departing from Under the premise of the principle of the invention, several improvements and modifications can also be made, these improvements and modifications are also considered as protection of the invention Within the scope of.The content being not described in detail in this specification belongs to the prior art well known to professional and technical personnel in the field.

Claims (10)

1. a kind of Craft Orbit design method, characterized in that it comprises:
Initial trajectory is designed and generated according to launch mission, obtains conditional parameter of entering the orbit；
Flight status parameter is calculated in the speed of launch inertial coordinate system and position according to carrier rocket；
Initial calculation condition is set, with last boosting first time Burnout altitude of the apogee, last second of shutdown point speed of boosting and Local trajectory tilt angle and target track height are constraint condition, according to initial calculation condition, enter the orbit conditional parameter and flight shape State parameter is calculated, and the deviation of constraint condition is obtained, and is optimized to initial calculation condition, and Optimal calculation condition is obtained, according to This is calculated and outputting standard ballistic data.
2. Craft Orbit design method as described in claim 1, it is characterised in that: the initial calculation condition includes end Boosting first time Burnout altitude of the apogee RE1, last boosting first time shutdown residual propellant quality DSMY41, last boosting second Secondary shutdown residual propellant quality DSMY42, end second of starting point locality trajectory tilt angle FIRE1 of boosting and last boosting first The attitude angle PHI4 of task.
3. Craft Orbit design method as claimed in claim 2, which is characterized in that carried out to initial calculation condition excellent Change, obtain Optimal calculation condition, specifically include:
Four adjustment thresholdings are set；
Last boosting shutdown residual propellant quality DSMY41 for the first time is adjusted, keeps last boosting first time Burnout altitude of the apogee inclined Absolute value of the difference is in the first thresholding；
Second of residual propellant quality DSMY42 that shuts down of last boosting is adjusted, the exhausted of second of shutdown point speed deviation of last boosting is made To value in the second thresholding；Also, second of starting point locality trajectory tilt angle FIRE1 of last boosting is adjusted, is made second of last boosting The absolute value of Burnout locality ballistic inclination deviation is in third thresholding；
Last boosting first time Burnout altitude of the apogee RE1 is adjusted, second of Burnout track major semiaxis deviation of last boosting is made Absolute value is in the 4th thresholding.
4. Craft Orbit design method as claimed in claim 3, it is characterised in that: the flight status parameter include into Rail point carrier rocket is real to the earth's core actual range RE, the practical absolute velocity V of injection point carrier rocket and injection point carrier rocket Border locality trajectory tilt angle η, and
Wherein, speed parameter of the carrier rocket in launch inertial coordinate system be VIxg, VIyg, VIzg, location parameter RExg, REyg,REzg。
5. Craft Orbit design method as claimed in claim 4, it is characterised in that: the conditional parameter of entering the orbit include into Rail point carrier rocket is to the earth's core theoretical distance bzRE, injection point carrier rocket theory absolute velocity VbzAnd injection point carrier rocket Theoretical locality trajectory tilt angle ηbz。
6. Craft Orbit design method as claimed in claim 5, it is characterised in that: the end boosting first time Burnout The calculation formula of altitude of the apogee RE1 is as follows:
RE1=bzRE-R0
Wherein, R0For earth radius.
7. Craft Orbit design method as claimed in claim 5, which is characterized in that the end boosting first time Burnout Altitude of the apogee deviation delta h1 are as follows:
Δ h1=h_apo-RE1
Wherein, h_apo is last boosting first time Burnout apogee computed altitude；
If the last inclined absolute value of the difference of boosting first time Burnout altitude of the apogee exceeds the first thresholding, last boosting is shut down for the first time The adjustment amount of residual propellant quality DSMY41 is Δ h/k1, k1For the first iteration coefficient.
8. Craft Orbit design method as claimed in claim 5, which is characterized in that described end second of Burnout of boosting Velocity deviation Δ V are as follows:
Δ V=V-Vbz
Described end second of Burnout locality ballistic inclination deviation Δ η of boosting are as follows:
Δ η=η-ηbz
If the last inclined absolute value of the difference of second of shutdown point speed of boosting exceeds the second thresholding of thresholding, second of remaining promote of shutting down The adjustment amount of agent quality DSMY42 is Δ V/k2, k2For secondary iteration coefficient；If last second of Burnout of boosting trajectory tilt angle it is inclined Absolute value of the difference exceeds third thresholding, then the adjustment amount of second of starting point locality trajectory tilt angle FIRE1 of last boosting is Δ η.
9. Craft Orbit design method as claimed in claim 5, which is characterized in that described end second of Burnout of boosting Track major semiaxis deviation delta h2 are as follows:
Δ h2=a-bzRE
Wherein, a is second of Burnout track major semiaxis of last boosting；
If last second of inclined absolute value of the difference of Burnout track major semiaxis of boosting exceeds the 4th thresholding, last boosting is shut down for the first time The adjustment amount of point altitude of the apogee RE1 is Δ h2.
10. Craft Orbit design method as claimed in claim 3, which is characterized in that the first thresholding e1For 50m, institute State the second thresholding e2For 0.01m/s, the third thresholding e3It is 0.001 °, the 4th thresholding e4For 100m.
CN201811321325.4A 2018-11-07 2018-11-07 Carrier rocket trajectory design method Active CN109583041B (en)

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CN112182772A (en) * 2020-10-11 2021-01-05 中国运载火箭技术研究院 Rocket propulsion control method, device and storage medium
CN112989559A (en) * 2021-01-19 2021-06-18 中国人民解放军63921部队 Analysis and evaluation method for contribution rate of liquid carrier rocket system
CN113342049A (en) * 2021-06-25 2021-09-03 北京航空航天大学 Multi-light source laser aircraft trajectory design method and system
CN114036780A (en) * 2021-12-06 2022-02-11 航天科工火箭技术有限公司 Rocket attitude angle design method constrained by space-based measurement and control
CN115618657A (en) * 2022-12-16 2023-01-17 中国人民解放军63921部队 Optimal design method for deployment task mode of medium-low orbit spacecraft
CN117113531A (en) * 2023-08-21 2023-11-24 江苏天兵航天科技有限公司 Method, system, equipment and storage medium for judging standby takeoff of carrier rocket

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CN107092731A (en) * 2017-04-05 2017-08-25 北京临近空间飞行器系统工程研究所 Specific thrust trajectory Integrated Optimization between sub-orbital booster level
CN108108524A (en) * 2017-11-30 2018-06-01 江西洪都航空工业集团有限责任公司 A kind of emulation mode of single-shot angle formula rocket assist transmitting unmanned plane takeoff phase

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Publication number Priority date Publication date Assignee Title
CN112182772A (en) * 2020-10-11 2021-01-05 中国运载火箭技术研究院 Rocket propulsion control method, device and storage medium
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CN112989559B (en) * 2021-01-19 2023-10-20 中国人民解放军63921部队 Analysis and evaluation method for contribution rate of liquid carrier rocket system
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CN115618657A (en) * 2022-12-16 2023-01-17 中国人民解放军63921部队 Optimal design method for deployment task mode of medium-low orbit spacecraft
CN117113531A (en) * 2023-08-21 2023-11-24 江苏天兵航天科技有限公司 Method, system, equipment and storage medium for judging standby takeoff of carrier rocket

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